Windage shield

ABSTRACT

A fan rotor module for a gas turbine engine, the fan rotor module comprising a drive arm, a fan disc and a windage shield attached at one or more points to a rear portion of the fan disc, wherein the fan rotor module comprises at least one balancing weight disposed at one or more of the points where the windage shield is attached to the rear portion of the fan disc.

CROSS-REFERENCE TO RELATED APPLICATIONS

This specification is based upon and claims the benefit of priority fromUnited Kingdom patent application number GB 1819251.8 filed on Nov. 27,2018, the entire contents of which are incorporated herein by reference.

BACKGROUND Field of the Disclosure

The present disclosure relates to a windage shield for mounting on arotor disc of a fan of a gas turbine engine, a fan of a gas turbineengine and a gas turbine engine.

Description of the Related Art

Gas turbine engines include a fan at the front of the engine. The fancomprises a plurality of radially extending blades mounted on a rotordisc which rotates about an axis. The fan is used as a first compressionstage for air passed to a combustor, and to direct air passing through aby-pass to provide thrust. Further compression stages for the airprovided to the combustor may also use rotating compressors, having aplurality of blades extending from a hub.

Rotation of a fan causes a pressure differential to be generated in aspace between an annulus filler and the rotor disc, with higher pressureair located at the position radially furthest from the axis of rotation,and lower pressure air located at the position radially closest to theaxis of rotation. Depending on the specific configuration of adjacenthardware in the engine, the air behind the fan may be stationary and somay have a lower pressure than the air under the annulus filler. At aradially outward position, the higher pressure air moves from under theannulus filler, into the low pressure region rearwards of the annulusfiller. To replace this air, new air is drawn into the area under theannulus filler, at a radially inward region. This sets up arecirculation which can reduce engine efficiency.

To prevent recirculation, it is known to provide a radially extendingshield, known as a windage shield. The windage shield is mounted on andextends from the rotor disc, and is disposed behind the rotor disc andthe annulus fillers. The windage shield may also include a lid thathelps form a smooth inner radial airflow surface and retain the annulusfiller against radial (centrifugal) forces.

SUMMARY

According to a first aspect there is provided a fan rotor module for agas turbine engine, the fan rotor module having a drive arm, a fan discand a windage shield attached at one or more points to a rear portion ofthe fan disc, wherein the fan rotor module comprises at least onebalancing weight disposed at one or more of the points where the windageshield is attached to the rear portion of the fan disc.

Having the balancing weight(s) disposed at one or more of the pointswhere the windage shield is attached to the rear portion of the fan discmay be convenient and/or may provide effective balancing of the fan discduring operation of the gas turbine engine.

One or more balancing weights may be disposed at each of the pointswhere the windage shield is attached to the rear portion of the fandisc.

The balancing weight(s) may each comprise one or more discrete bodies.

Each discrete body may have a known mass. Accordingly, the balancingweight(s) may be varied by varying the number of discrete bodies.

Harmful out-of-balance vibrational effects can be caused by an unevendistribution of mass about an axis of rotation of the fan disc (i.e. animbalanced fan). Imbalance of the fan rotor module and/or of the fandisc during operation of the gas turbine engine may be corrected,reduced or minimised by appropriate selection of the balancingweight(s). For a given gas turbine engine, the appropriate selection ofthe balancing weight(s) may depend on, for example, the size of the fan,the size of the fan disc and/or the radial location (i.e., the distancefrom the rotational axis of the fan) of the points where the windageshield is attached to the rear portion of the fan disc. Appropriateselection of the balancing weight(s) may also depend on, for example,scatter in fan blade masses (i.e., variation between masses of fanblades attached to the fan disc). Fan blade mass scatter can affect fanblades comprising any material, and can, for example, be prevalent infan blades comprising composite materials (e.g., carbon-fibre compositematerials).

The windage shield may be attached to the rear portion of the fan discby at least one mechanical fixing means. One or more of the mechanicalfixing means may comprise a nut and a bolt. The balancing weight(s) maycomprise one or more washers, e.g. cup washers, disposed on the bolt(s).

The drive arm may be a forward facing drive arm, or a rearward facingdrive arm.

Applying one or more balancing weights at one or more of the pointswhere the windage shield is attached to the rear portion of the fan discmakes use of existing attachment points to attach the balancing weightsto the fan disc. This arrangement may be particularly effective wherethere is no convenient existing location for attaching balancing weightsfor rear plane balancing of a fan disc. This arrangement may beeffective for fan rotor modules comprising a forward facing drive arm.

Disposing balancing weights at one or more points where the windageshield is attached to the rear portion of the fan disc may also meanthat the balancing weights are located further radially outward than ina conventional arrangement of a fan rotor module. A greater radialdistance between the balancing weights and the rotational axis of thefan rotor module may also mean less balancing weight is required toprovide the same balancing effect as a conventional arrangement (inwhich there is a lower radial distance between the balancing weights andthe rotational axis of the fan rotor module).

A second aspect provides a gas turbine engine comprising a fan rotormodule according to the first aspect.

The gas turbine engine may be an aircraft gas turbine engine.

A third aspect provides an aircraft comprising a gas turbine engineaccording to the second aspect.

A fourth aspect provides a method of rear-plane balancing of a fan rotormodule for a gas turbine engine, the fan rotor module having a drivearm, a fan disc and a windage shield attached at one or more points to arear portion of the fan disc, the method comprising:

applying at least one balancing weight at one or more of the pointswhere the windage shield is attached to the rear portion of the fandisc.

Applying the balancing weight(s) may comprise applying one or morebalancing weights at each of the points where the windage shield isattached to the rear portion of the fan disc.

The balancing weight(s) may each comprise one or more discrete bodies.Each discrete body may have a known mass. Accordingly, the balancingweight(s) may be varied by varying the number of discrete bodies.

The windage shield may be attached to the rear portion of the fan discby at least one mechanical fixing means. One or more of the mechanicalfixing means may comprise a nut and a bolt. The balancing weight(s) maycomprise one or more washers, e.g. cup washers, disposed on the bolts.

The drive arm may be a forward facing drive arm or a rearward facingdrive arm.

Applying one or more balancing weights at one or more of the pointswhere the windage shield is attached to the rear portion of the fan discmakes use of existing attachment points to attach the balancing weightsto the fan disc. This arrangement may be particularly effective wherethere is no convenient existing location for attaching balancing weightsfor rear plane balancing of a fan disc. This arrangement may beeffective for fan rotor modules comprising a forward facing drive arm.

Disposing balancing weights at one or more points where the windageshield is attached to the rear portion of the fan disc may also meanthat the balancing weights are located further radially outward than ina conventional arrangement of a fan rotor module. A greater radialdistance between the balancing weights and the rotational axis of thefan rotor module may also mean less balancing weight is required toprovide the same balancing effect as a conventional arrangement (inwhich there is a lower radial distance between the balancing weights andthe rotational axis of the fan rotor module).

As noted elsewhere herein, the present disclosure may relate to a gasturbine engine. Such a gas turbine engine may comprise an engine corecomprising a turbine, a combustor, a compressor, and a core shaftconnecting the turbine to the compressor. Such a gas turbine engine maycomprise a fan (having fan blades) located upstream of the engine core.

Arrangements of the present disclosure may be particularly, although notexclusively, beneficial for fans that are driven via a gearbox.Accordingly, the gas turbine engine may comprise a gearbox that receivesan input from the core shaft and outputs drive to the fan so as to drivethe fan at a lower rotational speed than the core shaft. The input tothe gearbox may be directly from the core shaft, or indirectly from thecore shaft, for example via a spur shaft and/or gear. The core shaft mayrigidly connect the turbine and the compressor, such that the turbineand compressor rotate at the same speed (with the fan rotating at alower speed).

The gas turbine engine as described and/or claimed herein may have anysuitable general architecture. For example, the gas turbine engine mayhave any desired number of shafts that connect turbines and compressors,for example one, two or three shafts. Purely by way of example, theturbine connected to the core shaft may be a first turbine, thecompressor connected to the core shaft may be a first compressor, andthe core shaft may be a first core shaft. The engine core may furthercomprise a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft may be arranged torotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axiallydownstream of the first compressor. The second compressor may bearranged to receive (for example directly receive, for example via agenerally annular duct) flow from the first compressor.

The gearbox may be arranged to be driven by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example the first core shaft in the example above). For example,the gearbox may be arranged to be driven only by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example only be the first core shaft, and not the second coreshaft, in the example above). Alternatively, the gearbox may be arrangedto be driven by any one or more shafts, for example the first and/orsecond shafts in the example above.

In any gas turbine engine as described and/or claimed herein, acombustor may be provided axially downstream of the fan andcompressor(s). For example, the combustor may be directly downstream of(for example at the exit of) the second compressor, where a secondcompressor is provided. By way of further example, the flow at the exitto the combustor may be provided to the inlet of the second turbine,where a second turbine is provided. The combustor may be providedupstream of the turbine(s).

The or each compressor (for example the first compressor and secondcompressor as described above) may comprise any number of stages, forexample multiple stages. Each stage may comprise a row of rotor bladesand a row of stator vanes, which may be variable stator vanes (in thattheir angle of incidence may be variable). The row of rotor blades andthe row of stator vanes may be axially offset from each other.

The or each turbine (for example the first turbine and second turbine asdescribed above) may comprise any number of stages, for example multiplestages. Each stage may comprise a row of rotor blades and a row ofstator vanes. The row of rotor blades and the row of stator vanes may beaxially offset from each other.

Each fan blade may be defined as having a radial span extending from aroot (or hub) at a radially inner gas-washed location, or 0% spanposition, to a tip at a 100% span position. The ratio of the radius ofthe fan blade at the hub to the radius of the fan blade at the tip maybe less than (or in the order of) any of: 0.4, 0.39, 0.38 0.37, 0.36,0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. Theratio of the radius of the fan blade at the hub to the radius of the fanblade at the tip may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds). These ratios may commonly be referred to as the hub-to-tipratio. The radius at the hub and the radius at the tip may both bemeasured at the leading edge (or axially forwardmost) part of the blade.The hub-to-tip ratio refers, of course, to the gas-washed portion of thefan blade, i.e. the portion radially outside any platform.

The radius of the fan may be measured between the engine centreline andthe tip of a fan blade at its leading edge. The fan diameter (which maysimply be twice the radius of the fan) may be greater than (or on theorder of) any of: 250 cm (around 100 inches), 260 cm, 270 cm (around 105inches), 280 cm (around 110 inches), 290 cm (around 115 inches), 300 cm(around 120 inches), 310 cm, 320 cm (around 125 inches), 330 cm (around130 inches), 340 cm (around 135 inches), 350 cm, 360 cm (around 140inches), 370 cm (around 145 inches), 380 (around 150 inches) cm or 390cm (around 155 inches). The fan diameter may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds).

The rotational speed of the fan may vary in use. Generally, therotational speed is lower for fans with a higher diameter. Purely by wayof non-limitative example, the rotational speed of the fan at cruiseconditions may be less than 2500 rpm, for example 2300 rpm. Purely byway of further non-limitative example, the rotational speed of the fanat cruise conditions for an engine having a fan diameter in the range offrom 250 cm to 300 cm (for example 250 cm to 280 cm) may be in the rangeof from 1700 rpm to 2500 rpm, for example in the range of from 1800 rpmto 2300 rpm, for example in the range of from 1900 rpm to 2100 rpm.Purely by way of further non-limitative example, the rotational speed ofthe fan at cruise conditions for an engine having a fan diameter in therange of from 320 cm to 380 cm may be in the range of from 1200 rpm to2000 rpm, for example in the range of from 1300 rpm to 1800 rpm, forexample in the range of from 1400 rpm to 1600 rpm.

In use of the gas turbine engine, the fan (with associated fan blades)rotates about a rotational axis. This rotation results in the tip of thefan blade moving with a velocity U_(tip). The work done by the fanblades 13 on the flow results in an enthalpy rise dH of the flow. A fantip loading may be defined as dH/U_(tip) ², where dH is the enthalpyrise (for example the 1-D average enthalpy rise) across the fan andU_(tip) is the (translational) velocity of the fan tip, for example atthe leading edge of the tip (which may be defined as fan tip radius atleading edge multiplied by angular speed). The fan tip loading at cruiseconditions may be greater than (or on the order of) any of: 0.3, 0.31,0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all units in thisparagraph being Jkg⁻¹K⁻¹/(ms⁻¹)²). The fan tip loading may be in aninclusive range bounded by any two of the values in the previoussentence (i.e. the values may form upper or lower bounds).

Gas turbine engines in accordance with the present disclosure may haveany desired bypass ratio, where the bypass ratio is defined as the ratioof the mass flow rate of the flow through the bypass duct to the massflow rate of the flow through the core at cruise conditions. In somearrangements the bypass ratio may be greater than (or on the order of)any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5,15, 15.5, 16, 16.5, or 17. The bypass ratio may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The bypass duct may besubstantially annular. The bypass duct may be radially outside theengine core. The radially outer surface of the bypass duct may bedefined by a nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/orclaimed herein may be defined as the ratio of the stagnation pressureupstream of the fan to the stagnation pressure at the exit of thehighest pressure compressor (before entry into the combustor). By way ofnon-limitative example, the overall pressure ratio of a gas turbineengine as described and/or claimed herein at cruise may be greater than(or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65,70, 75. The overall pressure ratio may be in an inclusive range boundedby any two of the values in the previous sentence (i.e. the values mayform upper or lower bounds).

Specific thrust of an engine may be defined as the net thrust of theengine divided by the total mass flow through the engine. At cruiseconditions, the specific thrust of an engine described and/or claimedherein may be less than (or on the order of) any of the following: 110Nkg⁻¹s, 105 Nkg⁻¹s, 100 Nkg⁻¹s, 95 Nkg⁻¹s, 90 Nkg⁻¹s, 85 Nkg⁻¹s or 80Nkg⁻¹s. The specific thrust may be in an inclusive range bounded by anytwo of the values in the previous sentence (i.e. the values may formupper or lower bounds). Such engines may be particularly efficient incomparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have anydesired maximum thrust. Purely by way of non-limitative example, a gasturbine as described and/or claimed herein may be capable of producing amaximum thrust of at least (or on the order of) any of the following:160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN,450 kN, 500 kN, or 550 kN. The maximum thrust may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds). The thrust referred to abovemay be the maximum net thrust at standard atmospheric conditions at sealevel plus 15 degrees C. (ambient pressure 101.3 kPa, temperature 30degrees C.), with the engine static.

In use, the temperature of the flow at the entry to the high pressureturbine may be particularly high. This temperature, which may bereferred to as TET, may be measured at the exit to the combustor, forexample immediately upstream of the first turbine vane, which itself maybe referred to as a nozzle guide vane. At cruise, the TET may be atleast (or on the order of) any of the following: 1400K, 1450K, 1500K,1550K, 1600K or 1650K. The TET at cruise may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The maximum TET in use of theengine may be, for example, at least (or on the order of) any of thefollowing: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. Themaximum TET may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds). The maximum TET may occur, for example, at a high thrustcondition, for example at a maximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/orclaimed herein may be manufactured from any suitable material orcombination of materials. For example at least a part of the fan bladeand/or aerofoil may be manufactured at least in part from a composite,for example a metal matrix composite and/or an organic matrix composite,such as carbon fibre. By way of further example at least a part of thefan blade and/or aerofoil may be manufactured at least in part from ametal, such as a titanium based metal or an aluminium based material(such as an aluminium-lithium alloy) or a steel based material. The fanblade may comprise at least two regions manufactured using differentmaterials. For example, the fan blade may have a protective leadingedge, which may be manufactured using a material that is better able toresist impact (for example from birds, ice or other material) than therest of the blade. Such a leading edge may, for example, be manufacturedusing titanium or a titanium-based alloy. Thus, purely by way ofexample, the fan blade may have a carbon-fibre or aluminium based body(such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion,from which the fan blades may extend, for example in a radial direction.The fan blades may be attached to the central portion in any desiredmanner. For example, each fan blade may comprise a fixture which mayengage a corresponding slot in the hub (or disc). Purely by way ofexample, such a fixture may be in the form of a dovetail that may slotinto and/or engage a corresponding slot in the hub/disc in order to fixthe fan blade to the hub/disc. By way of further example, the fan bladesmaybe formed integrally with a central portion. Such an arrangement maybe referred to as a blisk or a bling. Any suitable method may be used tomanufacture such a blisk or bling. For example, at least a part of thefan blades may be machined from a block and/or at least part of the fanblades may be attached to the hub/disc by welding, such as linearfriction welding.

The gas turbine engines described and/or claimed herein may or may notbe provided with a variable area nozzle (VAN). Such a variable areanozzle may allow the exit area of the bypass duct to be varied in use.The general principles of the present disclosure may apply to engineswith or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have anydesired number of fan blades, for example 16, 18, 20, or 22 fan blades.

As used herein, cruise conditions may mean cruise conditions of anaircraft to which the gas turbine engine is attached. Such cruiseconditions may be conventionally defined as the conditions atmid-cruise, for example the conditions experienced by the aircraftand/or engine at the midpoint (in terms of time and/or distance) betweentop of climb and start of descent.

Purely by way of example, the forward speed at the cruise condition maybe any point in the range of from Mach 0.7 to 0.9, for example 0.75 to0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Anysingle speed within these ranges may be the cruise condition. For someaircraft, the cruise conditions may be outside these ranges, for examplebelow Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond tostandard atmospheric conditions at an altitude that is in the range offrom 10000 m to 15000 m, for example in the range of from 10000 m to12000 m, for example in the range of from 10400 m to 11600 m (around38000 ft), for example in the range of from 10500 m to 11500 m, forexample in the range of from 10600 m to 11400 m, for example in therange of from 10700 m (around 35000 ft) to 11300 m, for example in therange of from 10800 m to 11200 m, for example in the range of from 10900m to 11100 m, for example on the order of 11000 m. The cruise conditionsmay correspond to standard atmospheric conditions at any given altitudein these ranges.

Purely by way of example, the cruise conditions may correspond to: aforward Mach number of 0.8; a pressure of 23000 Pa; and a temperature of−55 degrees C.

As used anywhere herein, “cruise” or “cruise conditions” may mean theaerodynamic design point. Such an aerodynamic design point (or ADP) maycorrespond to the conditions (comprising, for example, one or more ofthe Mach Number, environmental conditions and thrust requirement) forwhich the fan is designed to operate. This may mean, for example, theconditions at which the fan (or gas turbine engine) is designed to haveoptimum efficiency.

In use, a gas turbine engine described and/or claimed herein may operateat the cruise conditions defined elsewhere herein. Such cruiseconditions may be determined by the cruise conditions (for example themid-cruise conditions) of an aircraft to which at least one (for example2 or 4) gas turbine engine may be mounted in order to provide propulsivethrust.

The skilled person will appreciate that except where mutually exclusive,a feature or parameter described in relation to any one of the aboveaspects may be applied to any other aspect. Furthermore, except wheremutually exclusive, any feature or parameter described herein may beapplied to any aspect and/or combined with any other feature orparameter described herein.

BRIEF DESCRIPTION OF THE DRAWINGS

Embodiments will now be described by way of example only, with referenceto the Figures, in which:

FIG. 1 is a sectional side view of a gas turbine engine;

FIG. 2 is a close up sectional side view of an upstream portion of a gasturbine engine;

FIG. 3 is a partially cut-away view of a gearbox for a gas turbineengine;

FIG. 4 shows an arrangement of a part of a gas turbine engine comprisinga fan rotor module;

FIG. 5 shows an example of a windage shield mounted on a fan disc; and

FIG. 6 shows another example of a windage shield mounted on a fan disc.

DETAILED DESCRIPTION OF THE DISCLOSURE

FIG. 1 illustrates a gas turbine engine 10 having a principal rotationalaxis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23that generates two airflows: a core airflow A and a bypass airflow B.The gas turbine engine 10 comprises a core 11 that receives the coreairflow A. The engine core 11 comprises, in axial flow series, a lowpressure compressor 14, a high-pressure compressor 15, combustionequipment 16, a high-pressure turbine 17, a low pressure turbine 19 anda core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. Thebypass airflow B flows through the bypass duct 22. The fan 23 isattached to and driven by the low pressure turbine 19 via a shaft 26 andan epicyclic gearbox 30.

In use, the core airflow A is accelerated and compressed by the lowpressure compressor 14 and directed into the high pressure compressor 15where further compression takes place. The compressed air exhausted fromthe high pressure compressor 15 is directed into the combustionequipment 16 where it is mixed with fuel and the mixture is combusted.The resultant hot combustion products then expand through, and therebydrive, the high pressure and low pressure turbines 17, 19 before beingexhausted through the core exhaust nozzle 20 to provide some propulsivethrust. The high pressure turbine 17 drives the high pressure compressor15 by a suitable interconnecting shaft 27. The fan 23 generally providesthe majority of the propulsive thrust. The epicyclic gearbox 30 is areduction gearbox.

An exemplary arrangement for a geared fan gas turbine engine 10 is shownin FIG. 2. The low pressure turbine 19 (see FIG. 1) drives the shaft 26,which is coupled to a sun wheel, or sun gear, 28 of the epicyclic geararrangement 30. Radially outwardly of the sun gear 28 and intermeshingtherewith is a plurality of planet gears 32 that are coupled together bya planet carrier 34. The planet carrier 34 constrains the planet gears32 to precess around the sun gear 28 in synchronicity whilst enablingeach planet gear 32 to rotate about its own axis. The planet carrier 34is coupled via linkages 36 to the fan 23 in order to drive its rotationabout the engine axis 9. Radially outwardly of the planet gears 32 andintermeshing therewith is an annulus or ring gear 38 that is coupled,via linkages 40, to a stationary supporting structure 24.

Note that the terms “low pressure turbine” and “low pressure compressor”as used herein may be taken to mean the lowest pressure turbine stagesand lowest pressure compressor stages (i.e. not including the fan 23)respectively and/or the turbine and compressor stages that are connectedtogether by the interconnecting shaft 26 with the lowest rotationalspeed in the engine (i.e. not including the gearbox output shaft thatdrives the fan 23). In some literature, the “low pressure turbine” and“low pressure compressor” referred to herein may alternatively be knownas the “intermediate pressure turbine” and “intermediate pressurecompressor”. Where such alternative nomenclature is used, the fan 23 maybe referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox 30 is shown by way of example in greater detail inFIG. 3. Each of the sun gear 28, planet gears 32 and ring gear 38comprise teeth about their periphery to intermesh with the other gears.However, for clarity only exemplary portions of the teeth areillustrated in FIG. 3. There are four planet gears 32 illustrated,although it will be apparent to the skilled reader that more or fewerplanet gears 32 may be provided within the scope of the claimedinvention. Practical applications of a planetary epicyclic gearbox 30generally comprise at least three planet gears 32.

The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3is of the planetary type, in that the planet carrier 34 is coupled to anoutput shaft via linkages 36, with the ring gear 38 fixed. However, anyother suitable type of epicyclic gearbox 30 may be used. By way offurther example, the epicyclic gearbox 30 may be a star arrangement, inwhich the planet carrier 34 is held fixed, with the ring (or annulus)gear 38 allowed to rotate. In such an arrangement the fan 23 is drivenby the ring gear 38. By way of further alternative example, the gearbox30 may be a differential gearbox in which the ring gear 38 and theplanet carrier 34 are both allowed to rotate.

It will be appreciated that the arrangement shown in FIGS. 2 and 3 is byway of example only, and various alternatives are within the scope ofthe present disclosure. Purely by way of example, any suitablearrangement may be used for locating the gearbox 30 in the engine 10and/or for connecting the gearbox 30 to the engine 10. By way of furtherexample, the connections (such as the linkages 36, 40 in the FIG. 2example) between the gearbox 30 and other parts of the engine 10 (suchas the input shaft 26, the output shaft and the fixed structure 24) mayhave any desired degree of stiffness or flexibility. By way of furtherexample, any suitable arrangement of the bearings between rotating andstationary parts of the engine (for example between the input and outputshafts from the gearbox and the fixed structures, such as the gearboxcasing) may be used, and the disclosure is not limited to the exemplaryarrangement of FIG. 2. For example, where the gearbox 30 has a stararrangement (described above), the skilled person would readilyunderstand that the arrangement of output and support linkages andbearing locations would typically be different to that shown by way ofexample in FIG. 2.

Accordingly, the present disclosure extends to a gas turbine enginehaving any arrangement of gearbox styles (for example star orplanetary), support structures, input and output shaft arrangement, andbearing locations.

Optionally, the gearbox may drive additional and/or alternativecomponents (e.g. the intermediate pressure compressor and/or a boostercompressor).

Other gas turbine engines to which the present disclosure may be appliedmay have alternative configurations. For example, such engines may havean alternative number of compressors and/or turbines and/or analternative number of interconnecting shafts. By way of further example,the gas turbine engine shown in FIG. 1 has a split flow nozzle 20, 18meaning that the flow through the bypass duct 22 has its own nozzle thatis separate to and radially outside the core exhaust nozzle 20. However,this is not limiting, and any aspect of the present disclosure may alsoapply to engines in which the flow through the bypass duct 22 and theflow through the core 11 are mixed, or combined, before (or upstream of)a single nozzle, which may be referred to as a mixed flow nozzle. One orboth nozzles (whether mixed or split flow) may have a fixed or variablearea. Whilst the described example relates to a turbofan engine, thedisclosure may apply, for example, to any type of gas turbine engine,such as an open rotor (in which the fan stage is not surrounded by anacelle) or turboprop engine, for example. In some arrangements, the gasturbine engine 10 may not comprise a gearbox 30.

The geometry of the gas turbine engine 10, and components thereof, isdefined by a conventional axis system, comprising an axial direction(which is aligned with the rotational axis 9), a radial direction (inthe bottom-to-top direction in FIG. 1), and a circumferential direction(perpendicular to the page in the FIG. 1 view). The axial, radial andcircumferential directions are mutually perpendicular.

FIG. 4 shows a conventional arrangement of a part of a gas turbineengine comprising a fan rotor module. The fan rotor module comprises afan disc 100, a plurality of fan blades 105, a plurality of annulusfillers 110, a windage shield 115, a support ring 120, and a drive arm125 connected to the fan disc 100 at a first end and to a curvic joint130 at a second end. The windage shield 115 is mounted on a rear portionof the fan disc 100, disposed axially rearward of the annulus fillers110. The windage shield 115 is connected to the fan disc 100 via a nut135 and a bolt 140. In the arrangement shown, the drive arm 125 isrearward facing. The fan disc 100 is balanced in two planes (i.e., afrontal plane of the fan disc 100 and a rear plane of the fan disc 100).The fan disc 100 is balanced in the frontal plane by a first balancingweight or weights attached to an inner diameter of the support ring 120,outlined by a dashed circle A. The fan disc 100 is balanced in the rearplane by a second balancing weight or weights attached to an innerdiameter of the curvic joint 130, outlined by a dashed circle B.Balancing weights (i.e. first and second balancing weights) are appliedat both forward and rearward locations (i.e., the frontal plane and therear plane) of the fan 23 (FIGS. 1 and 2) of the gas turbine engine 10(FIGS. 1 and 2) to evenly distribute the mass of the fan rotor module(including the fan disc 100, fan blades 105, annulus fillers 110,windage shield 115, support ring 120 and other relevant componentsassociated with the fan rotor module) about the axis of rotation 9(FIGS. 1 and 2) of the fan disc 100, commonly referred to as “balancing”the fan rotor module. The fan rotor module is balanced to correct,reduce or minimise any harmful out-of-balance vibrational effects whichmay be caused by an uneven distribution of mass (i.e., an imbalanced fanrotor module) about the axis of rotation 9 of the fan disc 100.

FIG. 5 shows a conventional arrangement of a windage shield 115′ mountedon a fan disc 100′. The windage shield 115′ is mounted on a rear portionof the fan disc 100′. A shield portion 115A′ of the windage shield 115′extends axially rearward and radially outward from the fan disc 100′.The fan disc 100′ comprises disc posts 145′ extending radially outward.Although not shown in FIG. 5, a space or gap is positioned betweenadjacent disc posts 145′. The role of the disc posts 145′ is tocorrectly locate fan blades in the spaces between adjacent pairs of discposts 145′ when fan blades are mounted on the fan disc 100′. Anattachment flange 150′ extends radially outward from the disc posts 145′at a rear portion of the fan disc 100′. The windage shield 115′ isconnected to the fan disc 100′ at the attachment flange 150′ using a nut135′ and a bolt 140′. A head 155′ of the bolt 140′ is counter sunk intoa rear surface of the windage shield 115′.

FIG. 6 shows a windage shield 115″ in accordance with an embodiment ofthe disclosure mounted on a fan disc 100″. The windage shield 115″ ismounted on a rear portion of the fan disc 100″. A shield portion 115A″of the windage shield 115″ extends axially rearward and radially outwardfrom the fan disc 100″. The fan disc 100″ comprises disc posts 145″. Anattachment flange 150″ extends radially outward from the disc posts 145″at a rear portion of the fan disc 100″.

The windage shield 115″ is connected to the fan disc 100″ at theattachment flange 150″ using a nut 135″ and a bolt 140″. A balancingweight in the form of a cup washer 160″ is located on the bolt 140″ andis held securely between the attachment flange 150″ and the nut 135″. Ahead 155″ of the bolt 140″ is counter sunk into a rear surface of thewindage shield 115″.

Attaching a balancing weight at the attachment flange 150″ where thewindage shield 115″ is attached to a rear portion of the fan disc 100″is convenient as doing so makes use of an existing attachment point in agas turbine engine. It is therefore not necessary to design andmanufacture or incorporate a new attachment point in the gas turbineengine simply for the purpose of attaching balancing weights to correct,reduce or minimise imbalance of the fan disc 100″. Attaching a balancingweight at one or more points where a windage shield is attached to arear portion of a fan disc may be effective for gas turbine engines witha forward facing drive arm where there is no convenient location forrear plane balancing of a fan disc.

Attaching a balancing weight at one or more points where a windageshield is attached to a rear portion of a fan disc means that thebalancing weight is also further radially outward of the axis ofrotation of the fan disc, in comparison to a conventional arrangement(such as that shown in FIG. 4) where a rear plane balancing weight islocated on an inner diameter of a curvic joint 130. An equivalentbalancing effect to that achieved in a conventional arrangement cantherefore be achieved using balancing weights of lower mass located at agreater radial distance from the rotational axis of a fan disc (incomparison to a conventional arrangement).

In alternative arrangements, one or more balancing weights may bedisposed at any location at which a windage shield is attached to a rearportion of a fan disc. For example, for arrangements in which a windageshield is attached to a rear portion of a fan disc at a plurality ofpoints, one or more balancing weights may be disposed at one or more ofthe plurality of points.

A windage shield may be attached to a rear portion of a fan disc by atleast one mechanical fixing means. In the embodiment of FIG. 6, themechanical fixing means comprises a nut 135″ and a bolt 140″. Inalternative arrangements, the mechanical fixing means may be one or morerivets. The windage shield may be attached to the rear portion of thefan disc by any suitable means.

In the embodiment shown in FIG. 6, the balancing weight is a cup washer160″, attached to the fan disc 100″ via the nut 135″ and the bolt 140″passing through the windage shield 115″ and through the attachmentflange 150″. In alternative arrangements, the balancing weight(s) maycomprise one or more different types of washers, e.g. cup washers ordisc washers. The balancing weight(s) may alternatively be provided byone or more weighted nuts fastened to the bolt.

More generally, the balancing weight(s) may be provided as one or morediscrete bodies having a known mass. The one or more discrete bodies maycomprise a shape or form suitable for simple and convenient attachmentvia a windage shield attachment point. For example, the balancingweight(s) need not comprise an aperture for location on a bolt securinga windage shield to a rear portion of a fan disc, but may instead beheld securely by the compressive force between an attachment flange anda nut fastened to the bolt.

The balancing weight(s) comprising one or more discrete bodies may allowthe mass of the balancing weight(s) to be varied by varying the numberof discrete bodies used. The appropriate selection of the balancingweight(s) may depend on a variety of factors, including for example thesize of the fan, the size of the fan disc and the radial location (i.e.,the distance from the rotational axis of the fan) of the points wherethe windage shield is attached to a rear portion of the fan disc. Byutilising one or more discrete bodies, the balancing weight(s) may bemore easily varied to provide the required balancing weight to correctimbalance of a fan disc for a specific fan rotor module of a gas turbineengine. The number of discrete bodies may be varied before or after awindage shield is attached to a rear portion of the fan disc.

A method of rear plane balancing of a fan rotor module for a gas turbineengine, the fan rotor module comprising a drive arm, a fan disc and awindage shield attached at one or more points to a rear portion of thefan disc, may comprise applying at least one balancing weight at one ormore points where the windage shield is attached to the rear portion ofthe fan disc.

By attaching balancing weights where the windage shield is attached tothe rear portion of the fan disc, existing attachment locations can bemade use of. The design and/or manufacture of the fan rotor module neednot be altered to include an additional attachment point solely for thepurpose of attaching a balancing weight for rear plane balancing of thefan disc. Furthermore, the balancing weights attached at one or more ofthe points where the windage shield is attached to the rear portion ofthe fan disc may mean the balancing weights are located further radiallyoutward than the balancing weights of a conventional arrangement (forexample, where the balancing weights are located on an inner diameter ofa curvic joint attached to a drive arm connecting the curvic joint tothe fan disc). The greater radial distance between the balancing weightsand the rotational axis of the fan rotor module means that balancingweights of a lower mass may be used to provide the same balancing effectas a conventional arrangement.

The method may further comprise applying one or more balancing weightsat each of the points where the windage shield is attached to the rearportion of the fan disc. The balancing weight(s) may comprise one ormore discrete bodies. The balancing weights may therefore be varied byvarying the number of discrete bodies in each balancing weight. Thebalancing weight required may depend on the specific arrangement of thefan rotor module. Being able to vary the balancing weight by utilising avariable number of discrete bodies may therefore enable rapid and/orsimple adjustment of the balancing weight required for the specificcircumstances.

The windage shield may be attached to the rear portion by at least onemechanical fixing means. One or more of the mechanical fixing means maycomprise a nut and a bolt. The balancing weight(s) may comprise one ormore washers disposed on the bolt(s). The washers may be cup washers, ormay be disc washers.

The drive arm of the fan rotor module may be a forward facing drive arm,or may be a rearward facing drive arm.

The method described above may be effective for fan rotor modulescomprising a forward facing drive arm, in which there is no convenientexisting location for rear plane balancing of the fan rotor module.

It will be understood that the invention is not limited to theembodiments above-described and various modifications and improvementscan be made without departing from the concepts described herein. Exceptwhere mutually exclusive, any of the features may be employed separatelyor in combination with any other features and the disclosure extends toand includes all combinations and sub-combinations of one or morefeatures described herein.

We claim:
 1. A fan rotor module for a gas turbine engine, the fan rotormodule comprising a drive arm, a fan disc and a windage shield attachedat one or more points to a rear portion of the fan disc, wherein the fanrotor module comprises at least one balancing weight disposed at one ormore of the points where the windage shield is attached to the rearportion of the fan disc.
 2. The fan rotor module of claim 1, wherein oneor more balancing weights are disposed at each of the points where thewindage shield is attached to the rear portion of the fan disc.
 3. Thefan rotor module of claim 1, wherein the balancing weights each compriseone or more discrete bodies.
 4. The fan rotor module of claim 3, whereineach discrete body has a known mass.
 5. The fan rotor module of claim 1,wherein the windage shield is attached to the rear portion of the fandisc by at least one mechanical fixing means.
 6. The fan rotor module ofclaim 5, wherein one or more of the mechanical fixing means comprises anut and a bolt.
 7. The fan rotor module of claim 6, wherein one or morebalancing weights each comprise one or more washers disposed on thebolt(s).
 8. The fan rotor module of claim 1, wherein the drive arm is aforward facing drive arm or a rearward facing drive arm.
 9. A gasturbine engine including the fan rotor module of claim
 1. 10. Anaircraft including the gas turbine engine of claim
 9. 11. A method ofrear-plane balancing of a fan rotor module for a gas turbine engine, thefan rotor module having a drive arm, a fan disc and a windage shieldattached at one or more points to a rear portion of the fan disc, themethod comprising: applying at least one balancing weight at one or moreof the points where the windage shield is attached to the rear portionof the fan disc.
 12. The method of claim 11, further comprising applyingone or more balancing weights at each of the points where the windageshield is attached to the rear portion of the fan disc.
 13. The methodof claim 11, wherein the balancing weights each comprise one or morediscrete bodies.
 14. The method of claim 13, wherein the balancingweights are varied by varying the number of discrete bodies.
 15. Themethod of claim 11, wherein the windage shield is attached to the rearportion of the fan disc by at least one mechanical fixing means.
 16. Themethod of claim 15, wherein one or more of the mechanical fixing meanscomprises a nut and a bolt.
 17. The method of claim 16, wherein one ormore balancing weights each comprise one or more washers disposed on thebolt(s).
 18. The method of claim 10, wherein the drive arm is a forwardfacing drive arm or a rearward facing drive arm.